Gas turbine engine airfoil

ABSTRACT

A component for a gas turbine engine includes a platform that has a radially inner side and a radially outer side. A root portion extends from the radially inner portion of the platform. An airfoil extends from the radially outer side of the platform. The airfoil includes a pressure side that extends between a leading edge and a trailing edge. A suction side extends between the leading edge and the trailing edge. A bowed tip portion extends perpendicular to a mid-camber line of the airfoil.

BACKGROUND

This disclosure relates to a gas turbine engine airfoil. Moreparticularly, this disclosure relates to a gas turbine engine airfoilhaving an improved design.

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustorsection where it is mixed with fuel and ignited to generate a high-speedexhaust gas flow. The high-speed exhaust gas flow expands through theturbine section to drive the compressor and the fan section. Thecompressor section typically includes low and high pressure compressors,and the turbine section includes low and high pressure turbines.

The shape of an airfoil designed for turbomachinery applications is animportant characteristic. It is often a result of multidisciplinaryconsiderations including aerodynamics, durability, structures andmanufacturability. However, recent advances in the design ofaerodynamically high-performing, high-pressure turbine blades,particularly at the tip, have caused increased difficulties in thedesign of blades.

SUMMARY

In one exemplary embodiment, a component for a gas turbine engineincludes a platform that has a radially inner side and a radially outerside. A root portion extends from the radially inner portion of theplatform. An airfoil extends from the radially outer side of theplatform. The airfoil includes a pressure side that extends between aleading edge and a trailing edge. A suction side extends between theleading edge and the trailing edge. A bowed tip portion extendsperpendicular to a mid-camber line of the airfoil.

In a further embodiment of any of the above, the bowed tip portionextends in a circumferential direction and an axial direction.

In a further embodiment of any of the above, the bowed tip portionbeings at between 75% and 85% of a span of the airfoil.

In a further embodiment of any of the above, the bowed tip portionbegins at 80% of the span of the airfoil.

In a further embodiment of any of the above, the bowed tip portion isbowed between 3 and 19.6 degrees perpendicular to the mid-camber line ina circumferential direction.

In a further embodiment of any of the above, the bowed surface is bowedbetween 3 and 19.6 degrees perpendicular to the mid-camber line in anaxial direction.

In a further embodiment of any of the above, the bowed surface is bowedby the same degree in both the circumferential direction and the axialdirection.

In a further embodiment of any of the above, the bowed tip portionfollows a curvilinear profile beginning at 3 degrees and increasing to19.6 degrees.

In a further embodiment of any of the above, a leading edge of the bowedtip portion includes a rounded contour to reduce vibration andoxidation.

In another exemplary embodiment, a gas turbine engine includes acompressor section and a turbine section. A circumferential array ofairfoils are located in one of the compressor section and the turbinesection. Each of the airfoils include a pressure side that extendsbetween a leading edge and a trailing edge. A suction side extendsbetween the leading edge and the trailing edge. A bowed tip portionextends perpendicular to a mid-camber line of the airfoil.

In a further embodiment of any of the above, the bowed tip portionextends in a circumferential direction and an axial direction.

In a further embodiment of any of the above, the bowed tip portionbeings at between 75% and 85% of a span of the airfoil.

In a further embodiment of any of the above, the bowed tip portionbegins at 80% of the span of the airfoil.

In a further embodiment of any of the above, the bowed tip portion isbowed between 2 and 15 degrees perpendicular to the mid-camber line in acircumferential direction.

In a further embodiment of any of the above, the bowed surface is bowedbetween 3 and 19.6 degrees perpendicular to the mid-camber line in anaxial direction.

In a further embodiment of any of the above, the bowed surface is bowedby the same degree in both the circumferential direction and the axialdirection.

In a further embodiment of any of the above, the bowed tip portionfollows a curvilinear profile beginning at 3 degrees and increasing to19.6 degrees.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example gas turbine engine according toa first non-limiting embodiment.

FIG. 2 is a schematic view of a section of the gas turbine engine ofFIG. 1, such as a turbine section.

FIG. 3 is a schematic view of a turbine blade.

FIG. 4 is a cross-sectional view through an airfoil according to thisdisclosure.

FIG. 5 is an aft perspective view of the airfoil according to thisdisclosure.

FIG. 6 is a forward perspective view of the airfoil according to thisdisclosure.

FIG. 7 is a suction side perspective view of the airfoil according tothis disclosure.

FIG. 8 is a pressure side perspective view of the airfoil according tothis disclosure.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, and also drives air along acore flow path C for compression and communication into the combustorsection 26 then expansion through the turbine section 28. Althoughdepicted as a two-spool turbofan gas turbine engine in the disclosednon-limiting embodiment, it should be understood that the conceptsdescribed herein are not limited to use with two-spool turbofans as theteachings may be applied to other types of turbine engines includingthree-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

Referring to FIG. 2, a cross-sectional view through a high pressureturbine section 54 is illustrated. In the example high pressure turbinesection 54, first and second arrays of circumferentially spaced fixedvanes 60, 62 are axially spaced apart from one another. A first stagearray of circumferentially spaced turbine blades 64, mounted to a rotordisk 68, is arranged axially between the first and second fixed vanearrays. A second stage array of circumferentially spaced turbine blades66 is arranged aft of the second array of fixed vanes 62. It should beunderstood that any number of stages may be used. Moreover, thedisclosed airfoil may be used in a compressor section, turbine sectionand/or fixed or rotating stages.

The turbine blades each include a tip 80 adjacent to a blade outer airseal 70 of a case structure 72, which provides an outer flow path. Thefirst and second stage arrays of turbine vanes and first and secondstage arrays of turbine blades are arranged within a core flow path Cand are operatively connected to a spool 32, for example.

Each blade 64 includes an inner platform 76 respectively defining innerflow path. The platform inner platform 76 supports an airfoil 78 thatextends in a radial direction R, as shown in FIG. 3. It should beunderstood that the turbine vanes 60, 62 may be discrete from oneanother or arranged in integrated clusters. The airfoil 78 includes aleading edge 82 and a trailing edge 84.

The airfoil 78 is provided between pressure side 94 (predominantlyconcave) and suction side (predominantly convex) 96 in an airfoilthickness direction (FIG. 4), which is generally perpendicular to achord-wise direction provided between the leading and trailing edges 82,84. Multiple turbine blades 64 are arranged in a circumferentiallyspaced apart manner in a circumferential direction Y (FIG. 4). Theairfoil 78 includes multiple film cooling holes 90, 92 respectivelyschematically illustrated on the leading edge 82 and the pressure side94 (FIG. 4).

The turbine blades 64 are constructed from a high strength, heatresistant material such as a nickel-based or cobalt-based superalloy, orof a high temperature, stress resistant ceramic or composite material.In cooled configurations, internal fluid passages and external coolingapertures provide for a combination of impingement and film cooling.Other cooling approaches may be used such as trip strips, pedestals orother convective cooling techniques. In addition, one or more thermalbarrier coatings, abrasion-resistant coatings or other protectivecoatings may be applied to the turbine vanes 62.

FIGS. 3 and 4 schematically illustrate an airfoil including pressure andsuction sides joined at leading and trailing edges 82, 84. An attachmentor root 74 supports the platform 76. The root 74 may include a fir treethat is received in a correspondingly shaped slot in the rotor disk 68,as is known. The airfoil 78 extends a span from a support, such as aninner platform 76 to an end, such as a tip 80 in a radial direction Rfrom a radially outer side of the platform 76. The 0% span and the 100%span positions, respectively, correspond to the radial airfoil positionsat the support and the end. The leading and trailing edges 82, 84 arespaced apart from one another and an axial chord b_(x) length (FIG. 4)extends in the axial direction X.

As shown in FIG. 3, a radially outer end of the air foil 78 includes abowed tip portion 100. The bowed tip portion 100 is bowed in a directionperpendicular to a mid-camber line 98 (FIG. 4) of the airfoil 78 suchthat the bowed tip portion 100 extends towards the suction side 96 in acircumferential or tangential direction and towards the trailing edge 82in an axial downstream direction. The geometry of the bowed tip portion100 results in a non-pointed or rounded bowed tip leading edge 102 thatreduces vulnerable of a leading edge of the bowed tip portion 100 todamaging vibrations and oxidation.

As shown in FIG. 3, the bowed tip portion 100 beings at 80% of the spanof the airfoil 78 and continues to the tip 80. In another example, thebowed tip portion 100 begins between 75% and 85% of the bowed tipportion.

As shown in FIGS. 5 and 6, the bowed tip portion 100 bends in acircumferential direction towards the suction side 96 at an angle α. Inthe illustrated embodiment, the angle α is between 2 and 15 degrees.Similarly, as shown in FIGS. 7 and 8, the bowed tip portion 100 bends inan axially downstream direction towards at an angle β. In theillustrated embodiment, the angle β is between 3 and 19.6 degrees.Moreover, in the illustrated embodiment, the bowed tip portion 100 isbowed by the same degree in both the circumferential direction and theaxially downstream direction. In another example embodiment, the bowedtip portion 100 follows a curvilinear profile beginning at 3 degrees andincreasing to 19.6 degrees at the tip 80.

It should also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom. Although particular step sequencesare shown, described, and claimed, it should be understood that stepsmay be performed in any order, separated or combined unless otherwiseindicated and will still benefit from the present invention.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. A component for a gas turbine engine comprising:a platform having a radially inner side and a radially outer side; aroot portion extending from the radially inner portion of the platform;and an airfoil extending from the radially outer side of the platform,the airfoil including: a pressure side extending between a leading edgeand a trailing edge; a suction side extending between the leading edgeand the trailing edge; and a bowed tip portion extending perpendicularto a mid-camber line of the airfoil.
 2. The component of claim 1,wherein the bowed tip portion extends in a circumferential direction andan axial direction.
 3. The component of claim 2, wherein the bowed tipportion beings at between 75% and 85% of a span of the airfoil.
 4. Thecomponent of claim 3, wherein the bowed tip portion begins at 80% of thespan of the airfoil.
 5. The component of claim 1, wherein said bowed tipportion is bowed between 3 and 19.6 degrees perpendicular to themid-camber line in a circumferential direction.
 6. The component ofclaim 5, wherein the bowed surface is bowed between 3 and 19.6 degreesperpendicular to the mid-camber line in an axial direction.
 7. Thecomponent of claim 6, wherein the bowed surface is bowed by the samedegree in both the circumferential direction and the axial direction. 8.The component of claim 6, wherein the bowed tip portion follows acurvilinear profile beginning at 3 degrees and increasing to 19.6degrees.
 9. The component of claim 6, wherein a leading edge of thebowed tip portion includes a rounded contour to reduce vibration andoxidation.
 10. A gas turbine engine comprising: a compressor section anda turbine section; and a circumferential array of airfoils located inone of the compressor section and the turbine section, wherein each ofthe airfoils include: a pressure side extending between a leading edgeand a trailing edge; a suction side extending between the leading edgeand the trailing edge; and a bowed tip portion extending perpendicularto a mid-camber line of the airfoil.
 11. The gas turbine engine of claim10, wherein the bowed tip portion extends in a circumferential directionand an axial direction.
 12. The gas turbine engine of claim 11, whereinthe bowed tip portion beings at between 75% and 85% of a span of theairfoil.
 13. The gas turbine engine of claim 12, wherein the bowed tipportion begins at 80% of the span of the airfoil.
 14. The gas turbineengine of claim 10, wherein said bowed tip portion is bowed between 2and 15 degrees perpendicular to the mid-camber line in a circumferentialdirection.
 15. The gas turbine engine of claim 14, wherein the bowedsurface is bowed between 3 and 19.6 degrees perpendicular to themid-camber line in an axial direction.
 16. The gas turbine engine ofclaim 15, wherein the bowed surface is bowed by the same degree in boththe circumferential direction and the axial direction.
 17. The gasturbine engine of claim 15, wherein the bowed tip portion follows acurvilinear profile beginning at 3 degrees and increasing to 19.6degrees.